Gas-Turbine Aircraft Engine With Structural Surface Cooler

ABSTRACT

The present invention relates to a gas-turbine aircraft engine with a core engine surrounded by a bypass duct, with a substantially annular cooling element being provided in the area of a radially inner wall of the bypass duct downstream of stator vanes of the bypass duct.

This invention relates to a gas-turbine aircraft engine with a core engine surrounded by a bypass duct. Stator vanes are arranged in the bypass duct which guide the flow in the bypass duct downstream of a fan.

It is known from the state of the art to arrange cooling elements, heat exchangers or oil coolers downstream of the stator vanes of the bypass duct. This has the disadvantage that this area of the inner structure, which forms a part of the casing of the core engine, cannot be removed for servicing purposes, since the cooling elements are arranged there. Furthermore, air intakes are provided in this area of the core engine casing to supply air from the bypass duct to the inner area of the core engine, to be used there for cooling purposes, for example for cooling the turbine blades. All this has the result that the solutions known from the state of the art require a relatively large installation space and in addition hinder accessibility to components of the core engine, for example pumps, generators, bearings or the like. Furthermore, there is the disadvantage that the known structures require a large number of components and have a high weight. The manufacturing and assembly costs are considerable, too.

A broad aspect of the present invention is to provide a gas-turbine aircraft engine, which enables effective cooling by means of heat exchangers while avoiding the disadvantages of the state of the art.

It is a particular object of the present invention to provide solution to the above problematics by a combination of the features of claim 1. Further advantageous embodiments of the present invention become apparent from the sub-claims.

It is thus provided in accordance with the invention that a substantially annular cooling element is arranged in the area of a radially inner wall of the bypass duct downstream of stator vanes of the bypass duct.

In accordance with the invention, the terms oil cooler, heat exchanger and cooling element are used synonymously. In any event, the cooling element is intended to cool oil from the oil circuit of the gas-turbine aircraft engine.

The solution in accordance with the invention thus creates the possibility of arranging, on a limited axial length and with a reduced axial installation space, a high-efficiency cooling element. Since the cooling element is, in accordance with the invention, arranged directly downstream of the stator vane arrangement, assembly of the cooling element in accordance with the invention can be achieved simply and inexpensively.

It has proven particularly advantageous when the cooling element, which is annular overall, is composed of individual segments. The latter can for example be half-segments, such that two such half-segments can be joined together to form an annular cooling element extending around the entire circumference.

The cooling element in accordance with the invention is, in a preferred embodiment of the invention, flange-mounted downstream and directly on the stator vane ring or on the root structure of the stator vane ring, and is hence held by the latter. Further mounting elements can thus be dispensed with.

In a particularly favourable embodiment of the invention, the cooling element is designed as a surface cooler. It thus includes a plurality of ribs arranged in the flow direction, which considerably increase the surface area of the cooling element and thus improve its efficiency, allowing the overall axial length of the cooling element to be kept low.

In a further favourable embodiment of the invention, it is provided that the stator vanes inside the bypass duct are connected to the cooling element for heat transmission. Part of the heat from the cooling element can thus be transferred by heat conductance to the stator vanes such that the latter additionally contribute to cooling.

The cooling element is, in accordance with the invention, preferably provided with a plurality of cooling ducts extending in the circumferential direction. The flow through the cooling element is thus in the circumferential direction. As a result, an effective flow can be achieved with the medium to be cooled, for example oil.

It may be advantageous, for an inexpensive and simple manufacture of the cooling element in accordance with the invention, when the latter is made as an extruded section and bent into its annular shape during or after extrusion. With an extruded section of this type, it is then possible to design mounting points for flange-mounting of the cooling element to the supporting structure of the stator vanes (stator vane ring). The cooling ribs on the radially outer side of the extruded section can then be produced by means of a high-speed milling, normal milling or grinding process.

The cooling element in accordance with the invention is thus integrated into the supporting structure of the stator vane ring and/or into the surface of the core engine casing.

A further advantage of the solution in accordance with the invention is that the forces acting on the cooling element can be minimized such that the cooling element can have small dimensions without any danger of it being damaged by the forces generated during operation.

A further advantage of the solution in accordance with the invention is that the installations needed for connecting the fluid lines to the cooling element can be designed short and effective.

In the following, the present invention is described in light of the accompanying drawings showing an exemplary embodiment. In the drawings,

FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention,

FIG. 2 shows a partially perspective view the cooling element in accordance with the present invention,

FIG. 3 shows a perspective view of the cooling element as illustrated in FIG. 2,

FIG. 4 shows a perspective partial view of the cooling element in accordance with the present invention,

FIG. 5 shows a front-side view of the cooling element in accordance with the present invention in the axial direction, and

FIG. 6 shows a view, by analogy with FIG. 4, of the basic material for the cooling element in accordance with the present invention.

The gas-turbine engine 10 in accordance with FIG. 1 is an example of a turbomachine where the invention can be used. The following however makes clear that the invention can also be used in other turbomachines. The engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11, a fan 12 rotating inside a casing, an intermediate-pressure compressor 13, a high-pressure compressor 14, combustion chambers 15, a high-pressure turbine 16, an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19, all of which being arranged about a central engine axis 1.

The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.

The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.

FIG. 2 shows in a perspective partial view a simplified representation of a compressor drum 31 of the core engine 10. Furthermore, stator vanes 31 (OGV's) of the bypass duct 29 are illustrated. The reference numeral 33 shows in a simplified representation a carrier ring 33 which for example can include blade roots (not shown) of the stator vanes 31. FIG. 2 further shows a cooling element 30 in accordance with the invention, illustrated as a half-ring for greater clarity and described in detail in the following.

The cooling element 30 is connected (screwed) directly to the carrier ring 33.

FIG. 3 shows a perspective view of the cooling element shown in FIG. 2, likewise in a semi-circular design. Here a plurality of mounting attachments 34 are shown in particular, using which the cooling element 30 is attached.

FIGS. 4 and 5 show in an enlarged representation the cooling element 30. The latter is designed overall as a closed ring and consists preferably of individual sectors, for example two sectors each extending over 180° of the circumference. It is however also possible in accordance with the invention to subdivide the cooling element 30 into several sectors, such that individual part-elements of the cooling element 30 are formed.

As seen in FIGS. 4 and 5, the cooling element is designed as a surface cooler and includes a plurality of cooling ribs 35 extending in the axial direction relative to the engine axis 1. A plurality of cooling ducts 36 running parallel to one another in the circumferential direction are formed in the interior of the cooling element, through which cooling ducts flows the medium to be cooled (preferably oil). At their front side the cooling ducts 36 are closed at the end area of the cooling element 30 by dummy plugs, bolts or the like or by welding. The cooling ducts can have any cross-section, for example round, oval, angled or winding. A cross-connection of the individual cooling ducts 36 can be provided by transverse holes 37. The transverse hole 37 communicates with at least one connecting hole 38, which can be connected to a fitting or connecting element for supplying or removing the fluid to be cooled.

FIG. 6 shows an extruded basic material for a cooling element 30. The latter can either be extruded in circular form or appropriately bent after extrusion. FIG. 6 shows that the basic material is then machined to produce the cooling ribs 35. The basic material furthermore has a connecting web 39 which is then machined to form the individual mounting attachments 34. The material not needed is removed to save weight. Furthermore, a thickened area 40 remains (see FIG. 4) for accommodating the transverse hole 37 and the connecting holes 38. The area of the basic material in between them is also removed by machining to achieve a weight reduction.

LIST OF REFERENCE NUMERALS

-   1 Engine axis -   10 Gas-turbine engine/core engine -   11 Air inlet -   12 Fan rotating inside the casing -   13 Intermediate-pressure compressor -   14 High-pressure compressor -   15 Combustion chambers -   16 High-pressure turbine -   17 Intermediate-pressure turbine -   18 Low-pressure turbine -   19 Exhaust nozzle -   20 Guide vanes -   21 Engine casing/wall -   22 Compressor rotor blades -   23 Stator vanes -   24 Turbine blades -   25 Compressor drum or disk -   26 Turbine rotor hub -   27 Exhaust cone -   28 Bypass duct -   39 Cooling element -   31 Stator vane of the bypass duct/stator vane ring -   32 Compressor drum -   33 Carrier ring -   34 Mounting attachment -   35 Cooling ribs -   36 Cooling duct -   37 Transverse hole -   38 Connecting hole -   39 Connecting web -   40 Thickened area 

1. Gas-turbine aircraft engine with a core engine surrounded by a bypass duct, with a substantially annular cooling element being provided in the area of a radially inner wall of the bypass duct downstream of stator vanes of the bypass duct.
 2. Gas-turbine aircraft engine in accordance with claim 1, characterized in that the cooling element is designed as a surface cooler.
 3. Gas-turbine aircraft engine in accordance with claim 1, characterized in that the cooling element is arranged downstream of a rear area of a radially inner flange of a stator vane ring.
 4. Gas-turbine aircraft engine in accordance with claim 3, characterized in that the stator vanes are connected to the cooling element for heat transmission.
 5. Gas-turbine aircraft engine in accordance with claim 1, characterized in that the cooling element is provided with cooling ducts extending in the circumferential direction.
 6. Gas-turbine aircraft engine in accordance with claim 1, characterized in that the cooling element is provided as an extruded section bent into shape.
 7. Gas-turbine aircraft engine in accordance with Claim 1, characterized in that the cooling element on its radially outer side is provided with cooling ribs. 